1 Field of the Invention
The present invention generally relates to small propulsion systems for maneuvering spacecraft and, more particularly, is concerned with an electrothermal arcjet thruster employing any one of several different features for improving performance.
2 Description of the Prior Art
As conventionally known, an electrothermal arcjet thruster converts electrical energy to thermal energy by heat transfer from an arc discharge to a flowing propellant and from thermal energy to directed kinetic energy by expansion of the heated propellant through a nozzle. For an explanation from an historical perspective of arcjet thruster construction and operation and the problems associated with this type of electrothermal propulsion, attention is directed to the following publications: "Arcjet Thruster for Space Propulsion" by L. E. Wallner and J. Czika, Jr., NASA Tech Note D-2868, June 1965; "The Arc Heated Thermal Jet Engine" by F. G. Penzig, AD 671501, Holloman Air Force Base, March 1966; and "Physics of Electric Propulsion" by R. G. Jahn, McGraw-Hill Book Company, 1968. Attention is also directed to U.S. Pat. No. 4,548,033 to G.L. Cann.
Most electrothermal arcjet thrusters have as common features an anode in the form of a nozzle body and a cathode in the form of a cylindrical rod with a conical tip. The nozzle body has an arc chamber defined by a constrictor in a rearward portion of the body and a nozzle in a forward portion thereof. The cathode rod is aligned on the longitudinal axis of the nozzle body with its conical tip extending into the upstream end of the arc chamber in spaced relation to the constrictor so as to define a gap therebetween.
An electric arc is first initiated between the cathode rod and the anode nozzle body at the entrance to the constrictor. The arc is then forced downstream through the constrictor by pressurized vortex-like flow of a propellant gas introduced into the arc chamber about the cathode rod. The arc stabilizes and attaches at the nozzle. The propellant gas is heated in the region of the constrictor and in the region of arc diffusion at the mouth of the nozzle downstream of the exit from the constrictor. The super heated gas is then exhausted out the nozzle to achieve thrust.
Historically, pure propellants, typically ammonia (NH.sub.3) or hydrogen (H.sub.2), have been used in electrothermal arcjet thrusters. More recently, hydrazine (N.sub.2 H.sub.4) has been used as the propellant in arcjet thrusters developed by the assignee of the present invention. Propellants such as ammonia and hydrazine are storable in space as a liquid without refrigeration, whereas cryogenic propellants such as hydrogen and helium are not. Specific impulse levels achievable with propellants readily storable in space (e.g. NH.sub.3, N.sub.2 H.sub.4) have been limited to 800-1000 lbf-sec/lbm (pounds of force-second per pounds of mass), substantially lower than typical values of up to 1,500 lbf.sec/lbm achievable with cryogenic propellants (e.g. H.sub.2 He).
However, the performance advantage of cryogenic propellants due primarily to their very low molecular weights are offset by these same characteristics which make them difficult and expensive to store in space in useful quantities. Nonetheless, it would be desirable to be able to improve thruster performance to a level approaching that achievable using cryogenic propellants without adopting the difficulties normally associated with such propellants.